The present invention relates generally to the field of methods and apparatuses for curing damage repair patches. The invention is more specifically related to method and apparatuses for curing patches on contoured or curved surfaces. For example, new Navy aircraft such as the F/A-18 and the AV-8 are designed with increased amounts of epoxy composite materials on the aircraft surfaces. Techniques and devices for repairing damages that occur to such surfaces have become longstanding problems. Present field repair procedures for localized damage of epoxy composite materials involve the application of a patch that is vacuum, heat cured by a small rubberized heat blanket. There are inherent limitations with this approach. The major limitations are difficulty in maintaining a uniform temperature when curing patches over aircraft structural members and difficulty in creating and maintaining a vacuum and heating a patch over cured aircraft surfaces.
The use of composite materials such as epoxy composites employed in military aircraft construction is increasing. Epoxy materials either in monolithic sheets or in bonded honeycomb make up approximately ten percent of the weight and forty percent of the surface area of the F/A-18 aircraft. FIGS. 1 through 4 are, respectively, illustrations of the Navy's F-18, F-14, AV-8B, and the F-15 aircraft. The stipled areas 12 indicated on those aircraft illustrations represent the areas where composite material is used in the construction of the respective aircraft. As can be seen in FIG. 1, for instance, the F/A-18 employs composite materials in much greater quantities than previously utilized. The F/A-18 utilizes composite materials on the inner wing, outer wing, vertical stabilizer, horizontal stabilizer, trailing edge flap, rudder, speedbrake, dorsal covers, landing gear strut doors, inboard and outboard landing gear wheel doors, avionics bay doors and miscellaneous access covers. It is noted that although the wing and stabilizer skins are bolted onto the aircraft, they are not removable and therefore repair of damage in these areas must be accomplished, in most cases, with the wing and stabilizer skins in place because special jigs are required for their replacement.
The capability to repair otherwise minor localized damage to the composite structures of aircraft is extremely important if the operational readiness of the aircraft is to be maintained. Since damage is inevitable during aircraft operation and during ground and hangar movement of the aircraft, repair capability is extremely important. The present repair procedures of epoxy composites briefly mentioned above requires the application of a patch. The patch may involve replacement of a honeycomb core material, the use of preimpregnated fiber, foam, thin sheets of titanium or graphite and adhesive. The patch also requires application of heat and vacuum to ensure bonding of the materials.
There are further problems in repairing composite areas on aircraft. Areas such as the vertical stabilizers have underlying aluminum ribs and spars which act as heat sinks during the application of heat in the patching process. The effect of these heat sinks is to diminish the efficiency of the patching process.
FIG. 5 is an illustration of a basic temperature profile for a satisfactory epoxy cure plotted in terms of temperature versus time. The profile shown is for a particular graphite epoxy but the general requirements of a controlled temperature rise, a set temperature maintained for a fixed period of time, and a controlled temperature drop during cool-down are similar for different types of epoxy and are, therefore, generally represented by the graph of FIG. 5. The controlled temperature rise in the range of 3.degree. F./min to 5.degree. F./min maximum is necessary to ensure that the volatiles in the epoxy are driven off uniformly and that voids do not form. Voids leave the epoxy mixture structurally deficient. Also with the addition of a catalyst to the epoxy, an exothermic reaction takes places with heat generation. A rapid temperature rise coupled with the exothermic heat may mean that the epoxy mixture's temperature may rise to as high as 500.degree. F. again creating structural deficiences. After the controlled temperature rise takes place, a set temperature mut be uniformly maintained for a period of time. It is during this time that the epoxy cures, and its molecular matrix structure becomes intertwined and substantial structural strength achieved. The maximum temperature and its duration are variable. Higher temperatures cause curing of mixtures in a shorter time. However, tempering of underlying structural aluminum may be a consideration of higher tempertures. A representative set temperature and duration is 300.degree. F. for four hours. During this curing time the temperature must be maintained within .+-.10.degree. F. or in some cases .+-.20.degree. F. to ensure adequate shear strength. Finally, a cool-down rate must be maintained. If the temperature is allowed to drop too rapidly, the result is that microcracks will form. Microcracks allow the absorption of moisture which causes swelling, lowering of the transition temperature, and degradation of the compressive strength at elevated temperatures. Current literature suggests that if the cool-down rate is limited to 1.degree. F./min, the formation of microcracks is insignificant.
There are several heat blankets and heat blanket controllers presently on the market for utilization in patch curing as described above. They, however, have significant deficiencies which limit their use. For instance, Grumman Aerospace Corp. has developed a heat blanket/controller system suitable for some repairs of F-14's FIG. 6 is an isometric view of the Grumman Aerospace Corp. heat blanket 14 and FIG. 7 is a cross-section taken along lines VII--VII of FIG. 6 of the Grumman Aerospace Corp. heat blanket. Vaccuum hose attachment 16 is utilized to create a vacuum under the blanket 14 which seals by means of outer rubber ribs 18. The system includes a single heating element and a thermocouple system. This blanket design has the advantage of being simple and rugged and may work acceptably well in small plane or homogeneous repairs. However, it will not work on curved surfaces or areas where there are structural members below the surface which act as heat sinks.